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This study discusses the experimental results of the ballistic impact on thick-layered carbon fibre-reinforced polymer (CFRP), glass-based FRP (GFRP), and Kevlar-based FRP (KFRP) composite plates. Ten 10 mm plates of each type were manufactured via the hand lay-up and vacuum bagging techniques to produce plates with a thickness of 2 mm each, which were bonded together using epoxy adhesive and hot pressed at 75°C to obtain thick layered plates with a total thickness of 20 mm each. The specimen plates were tested according to the NIJ 0108.01 standard in a special firing room using a 9 mm G2 Elite handgun with a full metal jacketed round nose 9 mm × 19 mm Mu1-TJ bullets moving at a velocity of approximately 358 m/s. The results indicated that no total penetration occurred on any of the plates. The bullets were destroyed by impact, and the penetration depths of the CFRP, GFRP, and KFRP plates were 11.10, 9.95, and 22.22%, respectively. The layered structures in these plates appear to have changed the penetration behaviour compared with previous studies. In this study, delamination was the primary failure mode.

In airplane structural design, creating a strong yet lightweight structure is crucial. During the preliminary design phase, structural sizing optimization is necessary to achieve an efficient and lightweight structure that meets safety standards. Composite materials offer a high strength-to-weight ratio, enabling a lighter structure without compromising strength. This work focuses on optimizing composite structure wings by considering laminae thicknesses as design variables and minimizing structure weight as the objective, using the gradient-based method. The optimization constraints include the Tsai-Hill criterion, buckling factor, and flutter speed to ensure the structure's safety against static load, buckling phenomena, and flutter phenomena. The optimization results indicate that the buckling factor is the most critical constraint, carrying the highest weight. The weight of the wing structure decreased by 53% and 79% after optimizing with static and flutter constraints, respectively. Despite minimal changes in weight after optimizing with the buckling constraint, the structure now meets the buckling safety criteria and is safe from buckling. [ABSTRACT FROM AUTHOR]

This paper presents research on the use of anisogrid lattice structures in fighter wing applications. While the anisogrid lattice structure has been widely used in spacecraft structures, its implementation in main aircraft structures is still limited. The study is aimed at investigating the feasibility of utilizing an anisogrid lattice structure in fighter wing design. The analysis and optimization focus on determining the optimal weight of the composite wing structure, considering static, buckling, and flutter failure constraints. Various lift distributions, including triangular, Schrenk, and constant, are applied to evaluate the structure's response to static failure caused by aerodynamic loads. The anisogrid structure design incorporates inclined lattice elements between ribs and spars, with spar arrangement in the wing box featuring an anisogrid configuration. The anisogrid lattice structure is expected to produce higher bending and torsional stiffness compared to conventional orthogonal structures, producing better flutter and buckling characteristics. The optimized wing structure successfully meets static, buckling, and flutter load requirements at speeds below 500 m/s. The study showcases triangular, Schrenk, and constant load distributions resulting in half-wing masses of 504, 571, and 707 kg, respectively. The results show that flutter and buckling loads are no longer the critical loads in wing structural design but static load. [ABSTRACT FROM AUTHOR]

Due to their anisotropic behavior, composite structures are weak in transverse direction loading. produces transverse cracks, which for a laminated composite, may lead to delamination and total failure. The transition from transverse crack to delamination failure is important and the subject of recent studies. In this paper, a simulation of transverse crack and its transition to delamination on cross-ply laminate was studied extensively using a cohesive element Finite Element Method (FEM). A pre-cracked [0/90] composite laminate made of bamboo was modeled using ABAQUS/CAE. The specimen was in a three-point bending configuration. Cohesive elements were inserted in the middle of the 90° layer and in the interface between the 0° and 90° layer to simulate transverse crack propagation and its transition to delamination. A load–displacement graph was extracted from the simulation and analyzed. As the loading was given to the specimen, stress occurred in the laminates, concentrating near the pre-cracked region. When the stress reached the tensile transverse strength of the bamboo, transverse crack propagation initiated, indicated by the failure of transverse cohesive elements. The crack then propagated towards the interface of the [0/90] laminates. Soon after the crack reached the interface, delamination propagated along the interface, represented by the failure of the longitudinal cohesive elements. The result of the numerical study in the form of load–displacement graph shows a consistent pattern compared with the data found in the literature. The graph showed a linear path as the load increased and the crack propagated until a point where there was a load-drop in the graph, which showed that the crack was unstable and propagated quickly before it turned into delamination between the 0o and 90° plies. [ABSTRACT FROM AUTHOR]